MAJOR PROJ (Repaired)

Prediction of Supersonic Axisymmetric
Air Intake Performance through
Numerical Simulation
A PROJECT WORK
Submitted in fulfillment of the award of Degree of Bachelor of
Technology in Aeronautical Engineering
Submittedby
M.Akhilesh (12951A21C2)
Ch. Bapi Raju (12951A21C4)
Y.Ganesh Babu (12951A21D0)
T.Rohith Kumar (12951A21D8)
Under the Supervision of
PROF. K. BHARADWAJAN
Department of Aeronautical Engineering
INSTITUTE OF AERONAUTICAL ENGINEERING
DUNDIGAL – 500 043, HYDERABAD, TELANGANA STATE
April, 2016

i
CERTIFICATE
This is to certify that the work embodies in this dissertation entitled ‘Prediction Of
Supersonic Axisymmetric Air Intake Performance Through Numerical
Simulation’ being submitted by M.AKHILESH (12951A21C2), CH.BAPIRAJU
(12951A21C4), Y.GANESHBABU (12951A21D0), T.ROHITH KUMAR
(12951A21D8) for partial fulfillment of the requirement for the award of
Bachelor of Technology in Aeronautical Engineering discipline to Institute of
Aeronautical Engineering, Dundigal, Hyderabad, Telangana State, during the
academic year 2014-2015 is a record of Bonafide piece of work, undertaken by
him/her in the supervision of the undersigned.
Forwarded by
Dr. P. K. Dash Dr. D. Govardhan
DeanAcademics Aeronautical Engineering
IARE, Hyderabad IARE, Hyderabad
Approved and Supervised by
(Prof. K.Bharadwajan)
Aeronautical Engineering

ii
DECLARATION
We M.AKHILESH, CH.BAPIRAJU, Y.GANESH BABU, T.ROHITH KUMAR’,
are students of Bachelor of Technology in Aeronautical Engineering’,
session: 2012 - 16, Institute of Aeronautical Engineering, Dundigal, Hyderabad,
Telangana State, hereby declare that the work presented in this project work
entitled ‘Prediction of Supersonic Axisymmetric Air Intake Performance
Through Numerical Simulation’ is the outcome of our own bonafide work and is
correct to the best of our knowledge and this work has been undertaken taking care
of engineering ethics. It contains no material previously published or written by
another person nor material which has been accepted for the award of any other
degree or diploma of the university or other institute of higher learning, except
where due acknowledgment has been made in the text.
M.Akhilesh (12951A21C2)
Ch. Bapi Raju (12951A21C4)
Y.GaneshBabu (12951A21D0)
T.Rohith Kumar (12951A21D8)

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DUNDIGAL-500043, HYDERABAD, TELANGANA STATE
ACKNOWLEDGEMENT
We earnestly take the responsibility to acknowledge following distinguished personalities who
graciously allowed us to carry out my our project work successfully.
We would like to express our sincere thanks to Dr. Amalesh Barai, Principal,
Dr.D.Govardhan, Head of the department, Aeronautical Engineering,
Institute of Aeronautical Engineering, Hyderabad for their consistent guidance
and encouragement which lead to the successful completion of the project work.
We express our gratitude to our college Institute of Aeronautical Engineering, Hyderabad for
providing means of attaining of my most cherished goals. We profoundly thank my project
guide Prof. K. BHARADWAJAN who has been an excellent guide and also a great source of
inspiration to my work in carrying out my industrial oriented major project.
The satisfaction and euphoria that accompany the successful completion of the task
would be great but incomplete without the mention of the people who made it possible with their
constant guidance and encouragement crowns all the effort with success. In this context we
would like to thank all other staff members of the department who have extended their timely
help and eased our work.

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ABSTRACT
To predict performance evaluation of an supersonic axisymmetric air inlet through
numericalsimulation.To increase the individual component efficiency by overcoming the
problems at intake. First, the dimensions and physical parameters of the required inlet are
taken. Then generate the axisymmetric air intake model using GAMBIT, applying
relevant boundary conditions (for optimum mach number in supersonic flight conditions),
quantification of total drag experienced, analyze the aerodynamic performance and
individual component efficiency.
Perform numerical computations using FLUENT and obtain dynamic flow field
characteristics like velocity, total pressure distribution, mass flow ratio, boundary layer
seperation. Validation of obtaned data frm numercal results with the experimental data
provided are the desired outcomes of the proposed project

v
TABLES OF CONTENTS
S.No. Description Page. No.
1 Chapter 1 – Introduction 1
1.1 Over view of the project 3
2 Chapter 2 – literature review 4
2.1 Supersonic inlets 5
2.2 Ramjet stationing 6
2.3 Types of supersonic inlets 7
2.4 Axisymmetric inlets 7
2.5 Two dimensional 8
2.6 Types of compressions 8
3 Chapter 3 – supersonic Axisymmetric inlet 10
3.1 Introduction 10
3.2 Geometry 11
3.3 Modes of operation 11
3.4 performance 14
3.5 Influencing parameters 15
3.6 Relevant technical issues 16
4 Chapter 4 – problem definition 19
4.1 Flow separation 19
4.2 Pressure recovery 20
4.3 Inlet performance 21
4.4 Boundary layer bleed 22
4.5 Bleed system functions 23

vi
S.No. Description Page. No.
5 Chapter 5 – Introduction 24
5.1 Gambit 24
5.2 Create Geometry In Gambit 25
5.3 Mesh geometry in gambit 29
5.4 Specify Boundary Types In GAMBIT 33
5.5 Fluent 35
5.6 Initial Settings 37
5.7 Boundary Conditions 39
5.8 Operating Conditions 39
5.9 Solution Methods 40
5.10 Solution Controls 41
5.11 Monitors 42
5.12 Solution Initializations 42
5.13 Run Calculations 43
5.14 Results and Comparisons 43
6 Chapter 6 – Conclusion 57
7 Chapter 6 – References 58

vii
LIST OF FIGURES
Fig. No. Description Page No.
2.1 Schematic Sketch of Air Breathing Engine 4
2.2 Geometry of Supersonic Inlet 5
2.3 Ramjet 6
2.4 T-S Diagram Of Ramjet 6
2.5 Axisymmetric Inlet and Its Parts 7
2.6 Rectangular Supersonic Inlet 8
2.7 External Compression Inlet 8
2.8 Internal Compression Inlet 9
2.9 Mixed Compression Inlet 9
3.1 Supersonic Axisymmetric 10
3.2 Critical Inlet Operation 12
3.3 Subcritical Inlet Operation 12
3.4 Supercritical Inlet Operation 13
3.5 Boundary Layer Separation with Zones 16
3.6 Oblique Shock Diffuser 16
3.7 Phases Of Buzzing 17
3.8 Supersonic Inlet Relevant Technical Issues 18
4.1 Boundary Layer Separation with Zones 19
4.2 Geometry and Operation Of The SR-71 Mixed Compression
Inlet
20
4.3 Intake Shock System with Intake Critical (Ideal) 20
4.4 Starting Of an Intake 21
4.5 Represents Scramjet Flow Path With Boundary Layer Bleed
And Corresponding Station Numbering
22
4.6 Spike Bleed Exit Louver 23
5.1 Gambit Topology 25
5.2 Create real vertex 26
5.4 Creating Edges 28

viii
5.5 Creating Faces 29
5.6 Mesh edges 30
5.7 Edge mesh 31
5.8 Complete edge mesh 31
5.9 Mesh faces 32
5.10 Face mesh 33
5.11 Specify boundary types 34
5.12 Export mesh file 35
5.13 Gambit- Fluent 36
5.14 Fluent launcher 37
5.15 General 37
5.16 Models 38
5.17 Materials 38
5.18 Pressure Far Field 39
5.19 Operating Conditions 39
5.20 Reference Values 40
5.21 Solution Methods 41
5.22 Solution Controls 41
5.23 Residual Monitors 42
5.24 Solution Initialization 42
5.25 Run Calculation 43
5.26 Scaled Residuals with Slot 43
5.27 Scaled Residuals without Slot 44
5.28(a) Temperature Contour with slot 44
5.28(b) Temperature Contour without slot 45
5.29(a) Static Temperature Contour with slot 45
5.29(b) Static Temperature Contour without slot 46
5.30(a) Static Pressure Contour with slot 46

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5.30(b) Static Pressure Contour without slot 47
5.31(a) Velocity Contour with slot 47
5.31(b) Velocity Contour without slot 48
5.32(a) Enthalpy Contour with slot 48
5.33(a) Density Contour with slot 49
5.33(b) Density Contour without slot 49
5.34(a) Plot of Total Pressure with Slot 50
5.34(b) Plot of Total Pressure without Slot 50
5.35(a) Plot of Static Pressure with Slot 51
5.35(b) Plot of Static Pressure without Slot 51
5.36(a) Plot of Static Temperature with Slot 52
5.36(b) Plot of Static Temperature without Slot 52
5.37(a) Plot of Total Temperature with Slot 53
5.37(b) Plot of Total Temperature without Slot 53
5.38(a) Plot of velocity with Slot 54
5.38(b) Plot of velocity without Slot 54
5.39(a) Plot of Enthalpy with Slot 55
5.39(b) Plot of Enthalpy without Slot 55
5.40(a) Plot of Density with Slot 56
5.40(b) Plot of Density without Slot 56

x
LIST OF SYMBOLS AND ACRONYMS
° - Degrees
m - Mach number
SR - Strategic recoinnance
Po,f - Stagnation pressure at engine
face
P0, ∞ - Stagnation pressure in free
stream
Re - Reynolds Number
mR - mass flow ratio,
AOA - Angle of Attack

xi
LIST OF TABLES
Table No. Description Page No.
3.1 Modes of operation 13
5.1 Position of vertex points 25

1
ReportOverview
Chapter 1
Project
Overview
Chapter 2
Literature survey
1. Supersonic Inlets
2. Axisymetric Inlets
3) Types of
compression
4) Modes of
operation
Chapter3
Problem
defination
1) Objectives.
2) Inlet performance.
3) Performance of
intakes

2
CHAPTER 1
Introduction
1.1 Project Overview
An inlet for a supersonic aircraft, on the other hand, has a relatively sharp lip. The inlet lip is
sharpened to minimize the performance losses from shock waves that occur during supersonic
flight. For a supersonic aircraft, the inlet must slow the flow down to subsonic speeds before the
air reaches the compressor. Some supersonic inlets, like the one at the upper right, use a central
cone to shock the flow down to subsonic speeds. Other inlets, like the one shown at the lower
left, use flat hinged plates to generate the compression shocks, with the resulting inlet geometry
having a rectangular cross section.
This variable geometry inlet is used on the F-14 and F-15 fighter aircraft. More exotic inlet
shapes are used on some aircraft for a variety of reasons.
An inlet must operate efficiently over the entire flight envelope of the aircraft. At very low
aircraft speeds, or when just sitting on the runway, free stream air is pulled into the engine by the
compressor. In England, inlets are called intakes, which is a more accurate description of their
function at low aircraft speeds.
At high speeds, a good inlet will allow the aircraft to maneuver to high angles of attack and
sideslip without disrupting flow to the compressor. Because the inlet is so important to overall
aircraft operation, it is usually designed and tested by the airframe company, not the engine
manufacturer. But because inlet operation is so important to engine performance, all engine
manufacturers also employ inlet aerodynamicists. The amount of disruption of the flow is
characterized by a numerical inlet distortion index. Different airframes use different indices, but
all of the indices are based on ratios of the local variation of pressure to the average pressure at
the compressor face.
The ratio of the average total pressure at the compressor face to the free stream total pressure is
called the total pressure recovery. Pressure recovery is another inlet performance index; the
higher the value, the better the inlet. If the airflow demanded by the engine is much less than the
airflow that can be captured by the inlet, then the difference in airflow is spilled around the inlet.
The airflow mismatch can produce spillage drag on the aircraft. The aim of project is to provide
literature survey in the relevant state-of-the-art.

3
Aim
Performance
Prediction
Problem
Identification
Design and
Modelling
2d Modelling
and Geometry Check
Analysis Preprocessin
g and Solver
Post Processing
Results
and Plots
OVERVIEWOF THE PROJECT

4
CHAPTER 2
LITERATURE SURVEY
Air-Breathing Inlets
Most modern passenger and military aircraft are powered by gas turbine engines, which are also
called jet engines. There are several different types of gas turbine engines, but all turbine engines
have some parts in common. All turbine engines have an inlet to bring free stream air into the
engine. The inlet sits upstream of the compressor and, while the inlet does no work on the
flow, inlet performance has a strong influence on engine net thrust. As shown in the figures
above, inlets come in a variety of shapes and sizes with the specifics usually dictated by the
speed of the aircraft. The inlet interchanges the organized kinetic and random thermal energies of
the gas in an essentially adiabatic process. The perfect inlet would thus correspond to an
isentropic process. The primary purpose of the inlet is to bring the air required by the engine
from free stream conditions with minimum total pressure loss.[1]
Fig.2.1 Schematic Sketch of Air breathing Engine [1].
The performance of an inlet is related to the following characteristics:
 High total pressure ratio
 Controllable flow matching of requirements
 Good uniformity of flow
 Low installation drag
 Good starting and stability
 Low signatures (acoustics, radar, etc.)
 Minimum weight and cost
 Reliability goals

5
2.1 Supersonic Inlets
In the contrary, an inlet for a supersonic aircraft has a relatively sharp lip. The inlet lip is
sharpened to minimize the performance losses from shock waves that occur during supersonic
flight. For a supersonic aircraft, the inlet must slow the flow down to subsonic speeds before the
air reaches the compressor. Some supersonic inlets, like the one at the upper right, use a central
cone to shock the flow down to subsonic speeds. Other inlets, like the one shown at the lower
left, use flat hinged plates to generate the compression shocks, with the resulting inlet geometry
having a rectangular cross section. This variable geometry inlet is used on the F-14 and F-15
fighter aircraft. More exotic inlet shapes are used on some aircraft for a variety of reasons. The
inlets of the Mach 3+ SR-71 aircraft are specially designed to allow cruising flight at high speed
[3].
Fig.2.2: Geometry of supersonic inlet [3].
2.2.1 Principle:
Flow in a supersonic inlet is much more problematic than the flow in a subsonic inlet. The big
difference is that flow moving faster than the speed of sound has no knowledge of what is ahead
of it. The result of this makes the design of a supersonic inlet similar to laying out the course for
a slalom skier, only in this case the skier is blind.

6
2.2 RAMJET Stationing:
Fig.2.3: Ramjet [Ref 4]
P-v DIAGRAM T-S DIAGRAM T-P DIAGRAM
Fig.2.4 T-S diagram of ramjet [Ref 3]

7
2.3 Types of Supersonic Inlets
 Axisymmetric or two-dimensional
• Axisymmetric: central cone for shock fixture
• Two-dimensional: rectangular cross-section
 Variable or fixed geometry
• Variable: the central cone may be movable or in a rectangular intake,
one of the walls may be adjustable
• Fixed: Geometry is fixed.
2.4 Axisymmetric Inlet
According to Bendot et al. (1984) the performance of an axisymmetric inlet deteriorates rapidly
with an increase in angles of attack for two-dimensional and chin inlet configurations an
improvement in performance with increasing positive (nose-up) angles of attack is encountered.
The role of engine inlet is very vital for all air-breathing propulsion systems; especially for high
supersonic in which all the compression process is performed by the inlet.
Fig.2.5 Axisymmetric Inlet and Its Parts [Ref1].

8
2.5 Two Dimensional Inlets:
Fig.2.6 Rectangular supersonic inlet [Ref 2].
2.6 Types of compressions:
2.6.1 External Compression:
The supersonic portion of the compression is done externally, ahead of the cowl lip. An
advantage of a two-wedge external compression inlet is that the construction is relatively
simple and the optimum configuration is normally determined experimentally in a wind-tunnel.
Fig.2.7 External compression inlet [Ref 6]

9
2.6.2 Internal Compression:
Supersonic compression is done internally aft of the cowl lip. This type of inlet has a low drag
value because external deflection and disturbance of the entering flow are prevented.
Fig.2.8 Internal Compression Inlet [Ref 6]
2.6.3 Mixed Compression:
Supersonic compression is done both externally, forward of the cowl lip and internally, aft of the
cowl lip. The advantage of a better stagnation pressure recovery is cancelled out to a great extent
by the need to control Bow separation in the inlet duct. The flow separation is caused by the
interaction of the internal
Shock systems (oblique and normal shocks) with the boundary layer.
Fig.2.9 Mixed compression inlet [Ref 6]

10
CHAPTER 3
SUPERSONIC AXISYMMETRIC INLET
3.1 Introduction
The supersonic inlet consists of a spike (center-body or fore-body) and an integrated duct. The
initial compression is done by the spike. When designing a supersonic inlet, an increase in the
flight Mach number requires the increase in the number of oblique shocks in order to save total
pressure recovery. Therefore, the principle of staging a supersonic compression like the
Oswatitsch principle [Ref 3] is used to reduce the inlet losses in a most efficient manner.
Fig.3.1 Supersonic Axisymmetric Inlet

11
3.2 Geometry
An inlet generally consists of the following components:
 Ramp: An intake ramp is a rectangular, plate-like device within the air intake of a jet
engine, designed to generate a shock wave to aid the inlet compression process
at supersonic speeds. The ramp sits at an acute angle to deflect the intake air from the
longitudinal direction
 Throat: A throat is minimum area between the centre body and cowl.
 Cowl: Cowl is upper and lower part of the engine.
 Subsonic diffuser: It is a divergent profile to increase the pressure of the flow. [3]
3.3 Modes Of Operation:
3.3.1 Critical Operation
The condition when the inlet can accept the mass flow of air required to position the terminal
shock just inside the cowl lip is called critical inlet operation.
Modes of
operation
Critical Subcritical Supercritical

12
Fig.3.2 Critical Inlet Operation [Ref 4].
3.3.2 Subcritical Operation
The condition when the inlet is not matched to the engine, due to which the normal shock moves
upstream and stays in front of cowl lip, is called as sub-critical operation
Fig.3.3 Subcritical Inlet Operation [Ref 4].

13
3.3.3 Supercritical Operation:
The condition when the inlet cannot capture the mass flow required by the engine and the
terminal shock is sucked into the diffuser is called super – critical operation
Fig.3.4 Supercritical inlet operation [Ref 4].
Modes of
operation
Axisymmetric inlet Sketch
Position of
Normal shock
wave
Sub-Critical
The normal shock
moves upstream
and stays in front
of cowl lip
Critical
The terminal
shock just inside
the cowl lip
Super-Critical
The terminal
shock is sucked
into the diffuser
3.1 Modes of operation

14
3.4 Performance:
The performance of an inlet is characterized by three important parameters, namely pressure
recovery, mass flow ratio and boundary layer bleed. Each one of the above-mentioned has an
effect on the overall or total drag figure of an inlet.
3.4.1 Pressure Recovery
The inlet total pressure recovery is defined by Seddon (1988:5) as the ratio of the mean total
pressure at the engine face to the total pressure available in the free stream, infinitely far
upstream of the inlet.
Po,f = Stagnation pressure at engine face
P0, ∞ = Stagnation pressure in free stream
The total pressure recovery is an indication of the maximum pressure available in the combustion
chamber and resultant thrust that can be developed. With an increase in free stream (flight) Mach
numbers, high shock losses and shock wave boundary layer interactions cause a decrease in total
pressure recovery. Methods to remedy these high shock losses, ego variable compression surface
where oblique shocks are prevented from entering into the inlet, can result in an increase in drag.
This is as a result of the greater angle through which the flow is turned before it enters the inlet
opening.
3.4.2 Mass Flow Ratio
The mass flow ratio is defined by Gregoriou (1985:7} as the ratio of air mass flow at inlet entry
to air mass flow at free stream conditions.
Where,

15
This mass flow ratio, mR, can also be expressed as an area ratio since the equation for mass
flow (Hall 1977:12) can be written as follows:
3.5 Influencing Parameters:
3.5.1 Total Pressure Ratio:
The ratio total pressure at inlet is given by the ratio of total pressure leaving at stage 2 and the
total pressure entering at stage 0.
Πd = Pt2/Pt0
3.5.2 Total Temperature Ratio:
The ratio total temperature at inlet is given by the ratio of total temperature leaving at stage 2
and the total temperature entering at stage 0.
Ʈd = Tt2/Tt0
3.5.3 Isentropic Diffuser Efficiency
3.5.4 Area Mach Relation

16
3.6 Problems
3.6.1 Boundary Layer Separation
 Separation of the external flow in zone 1 may result from local high velocities and
subsequent deceleration over the outer surface and it leads to high nacelle drag.
 Separation on the internal surfaces may take place in either zone 2 or zone 3, depending
on the geometry of the duct and the operating conditions.
 Zone 3 may be the scene of quite large adverse pressure gradients since the flow
accelerates around the nose of the center body and then decelerates as the curvature
decreases.
Fig.3.5 Boundary Layer Separation with Zones [Ref 7]
3.6.2 External Deceleration
The simplest and most practical external deceleration in simple oblique shock wave is in some
cases, a series of oblique shock waves.
Fig.3.6 Oblique shock diffuser [Ref 7]

17
3.6.3 Buzzing:
Buzz is an airflow instability caused by the shock waves rapidly being alternately swallowed and
expelled at the inlet of the duct and occurs in supersonic intakes at subcritical operations. It starts
when the aircraft begins to fly at or near the speed of sound. At these speeds sonic shock waves
are developed that if not controlled will give high duct loss in pressure and airflow and will set
up vibrating conditions in the inlet duct, called inlet Buzz.[8].
Fig.3.7 Phases of buzzing [Ref 8].

18
Fig.3.8 Supersonic Inlet Relevant Technical Issues
We will focus on the problem spike bleed and we will proceed to how to solve it efficiently.
Supersonic
inlet
relevant
technical
issues
Spike bleed
Buzzing
External deceleration
Unstart
Region of
interest

19
CHAPTER 4
Problem Definition
The boundary layer on the cone is stretched as it moves up the cone preventing flow separation ,
but for the internal compression and the subsonic compression the boundary layer still tends to
separate and usually is sucked through tiny holes in the wall. As a side note on the aero spike
engine the boundary layer gets thicker towards the end of the cone as needed for the greater
speed difference between the air molecules just on the surface of the cone and the fully
accelerated stream of air.
Fig.4.1 Boundary Layer Separation with Zones [Ref 7]
OBJECTIVES
4.1 FLOW SEPERATION
Supersonic flow over spiked cylinders nonsteady regimes
can occur in which a separation zone is periodically
generated at the spike. The physical pattern of flow with
separation zone fluctuations and have determined the boundaries of existence of the nonsteady
regime as a function ratio between the spike length and diameter of the cylinder.

20
Fig.4.2 Details of the Geometry and Operation of the SR-71 Mixed Compression Inlet
By J. Thomas Anderson Technical Fellow Emeritus Lockheed Martin [6].
4.2 Pressure Recovery
Pressure recovery charecteristics of conical spike inlet with a fixed area bypass located in the top
or bottom of the diffuser are presented for fligh mach numbers of 1.6, 1.8, and 2.0 for angle of
attack 0 degrees to 9 degrees.Top or bottom location of bypass did not have significant effects on
diffuser pressure recovery,bypass mass flow ratio,or drag coefficient over the range of angle of
attack,flight mach numbers and stable engine mass flow are investigated.
Fig.4.3 intake shock system with intake critical
At a flight Mach number of 2.0 the discharge of 14 percent of the critical mass flow of the inlet
by means of by pass increased the drag only one-fifth of the additive drag that would result for
equivalent spillage behind an inlet normal shock without significant reductions in pressure
recovery.

21
4.3 Inlet Performance
Supersonic diffusers are characterized by the presence of shocks. However before the intake
operates in a supersonic flow, it must pass through the subsonic flow regime. In some types of
supersonic intakes, establishing a shock system with minimal losses is not easy. The process of
establishing a stable shock system is referred to as Starting of an intake.
Fig.4.4 Starting of an Intake
External compression intakes complete the supersonic diffusion outside the covered portion of
the intake .These intakes usually have one or more oblique shocks followed by a normal shock.
Depending upon the location of these shocks, the intake may operate in subcritical, critical or
supercritical modes.
4.3 Performance of Intakes
o Performance parameters
o Sources of losses
o Starting problem in supersonic intakes.
o Modes of operation of an external compression intake

22
4.4 Boundary Layer Bleed
The formation of a boundary layer on the compression surfaces of the inlet and the interaction
Of shocks with the boundary layer cause detrimental pressure gradients and the flow separates
from the diffuser walls (Seddon 1988). It will result in an unsteady non-uniform distribution
off low at the engine face, a loss in stagnation pressure recovery, increased internal drag and
too high Mach numbers due to the reduced flow area at the separation region. This is one of
The most difficult problems to solve when designing supersonic air-inlet systems
Fig.4.5 Represents Scramjet Flow Path With Boundary Layer Bleed And Corresponding
Station Numbering [Ref 5].
Spike bleed air flows through the bleed slot into the translating spike and then enters the front of
the fixed centerbody. The air continues through the fixed centerbody, out the four struts, and then
flows overboard through louvers. This flow path is illustrated in Figure. [8]

23
Fig.4.6 Spike Bleed Exit Louver[8]
4.5 BleedSystem Functions:
4.5.1 Purpose
o Mass flow rate control
o Cabin air quality
o Cooling air systems
o Shock Wave control
Location
o Before Compressor-Supersonic aircraft
Current problems
o Mechanically complex
o Bypass airflow wasted

24
Chapter 5
Design and Analysis
To solve any fluid flow or heat transfer problem, one needs to rely on CFD software
FLUENT+GAMBIT. It serves various options like conduction, convection, heating, cooling and
fluid-structure interaction (FSI) problem. Recently FLUENT is now takeover by ANSYS and
named to it as ANSYS CFX. Though people who have already bought FLUENT and GAMBIT
separately can use following procedure to analyze a given fluid flow and heat transfer problem
via combination of these two software.
GAMBIT and FLUENT are tools to analysis the fluid flow problems and the branch of science
for this problem is known as Computation Fluid Dynamics (CFD)
5.1 GAMBIT
 GAMBIT full form is Geometry And Mesh Building Intelligent Tool
 GAMBIT is used for pre-processing operation (which is required before starting of
solution) of fluid flow problem which includes following operations
 Geometry creation (specifies the domain of fluid flow problem). It can even model 1D,
2D and 3D domain
 Mesh generation (discretization of domain to solve governing equations at each cell)
which allows solving every governing equation at each node created in mesh.
 Specifying the boundary zones (name & type) to apply boundary conditions for
problem. One need to be very careful in choosing boundary conditions as this will decide
the nature and behavior of any physical phenomena associated with the problem at hand.
 GAMBIT export the file containing all the data related to pre-processing.
GAMBIT is a type of modeling software and out of it will serve as input to the FLUENT.

25
Fig.5.1 Gambit Topology
5.2 Create Geometryin GAMBIT
5.2.1 Create Vertices
The coordinates needed for the mesh are shown below
VERTEX X Y Z
A 0 0 0
B 2.40 0.756 0
C 3.50 1.311 0
D 3.70 1.351 0
E 3.80 1.367 0
F 3.90 1.390 0
G 7.00 1.390 0
H 8.00 1.362 0
I 8.75 1.370 0
J 9.12 1.348 0
K 9.50 1.320 0
L 10.14 0 0
M 10.14 1.474 0
N 3.50 1.974 0
O 10.14 1.974 0
P 10.14 2.421 0
Table 5.1: Position of vertex points

26
Fig.5.2 Create real vertex
Using bottom up approach, we start by creating vertices of the geometry using the coordinate
given.
Operation Toolpad > Geometry Command Button > Vertex Command Button >
Create Vertex
Create the vertices by entering the coordinates under Global and the label under Label:

27
Click the FIT TO WINDOW button to scale the display so that you can see all the vertices.
The resulting image should look like this:
Fig.5.3 Vertex points
5.2.2 Create Edges
Now we can create the edges using the vertices created.
Operation Toolpad > Geometry Command Button > Edge Command Button >
Create Edge
Create the edge AB by selecting the vertex A followed by vertex B. Enter AB for Label.
Click Apply. GAMBIT will create the edge.
Similarly, create the edges BC, CD, DE, EF, FA , etc., Click on the to select the vertices
from the list and move them to the picked list. You can also hold the shift button and mouse
click the vertices for selection. The resulting image should look like this.

28
Fig.5.4 Creating Edges
5.2.3 Create Faces
The edges we have created can be joined together to form faces. We will need to define
seven faces.
Operation Toolpad > Geometry Command Button > Face Command Button >
Form Face
This brings up the Create Face From Wireframe menu. Recall that we had selected vertices
in order to create edges. Similarly, we will select edges in order to form a face.
To create the face1, select the edges AB, BC, CE, EF and FA. Enter face1 for the label and
click Apply. GAMBIT will tell you that it has "Created face: face1'' in the transcript window.
Similarly, create the face face2 by selecting GH, HI, and GI.

29
Fig.5.5 Creating Faces
5.3 MeshGeometry in GAMBIT
Operation Toolpad > MeshCommand Button > Edge Command Button ** *>
Mesh Edges*
Select the edge AB. The edge will change color and an arrow will appear on the edge. This
indicates that you are ready to mesh this edge. FIRST LENGTH, Select interval size under
Spacing. Enter 0.01 for interval

30
Fig.5.6 Mesh edges
Do the same for edge CD and CF.

31
Fig.5.7 Edge mesh
Fig.5.8 Complete edge mesh

32
Now that the appropriate edge meshes have been specified, mesh the face face1:
Operation Toolpad > MeshCommand Button > Face Command Button > Mesh
Faces
Select the face1. The face will change color. You can use the defaults of Tri (i.e. triangles)
and Pave. Click Apply.
Fig.5.9 Mesh faces

33
The meshed face should look as follows:
Fig.5.10 Face mesh
Next for meshing face2:
 Operation Toolpad > MeshCommand Button > Edge Command
Button> mesh nodes
 Operation Toolpad > MeshCommand Button > Edge Command
Button>Link command button
First select edge CF and then select DE click
Apply Second select edge EF and then select CD
click Apply
5.4 Specify Boundary Types in GAMBIT
We'll label the boundary JKLM as pressure far-field, MN pressure inlet, JI,FE as pressure
oulet, ABCDE, GHIF as wall and NA as symmetry. Recall that these will be the names that
show up
under boundary zones when the mesh is read into FLUENT.

34
5.4.1 Define Boundary Types
Operation Toolpad > Zones Command Button > Specify Boundary Types
Fig.5.11 Specify boundary types
Save Your Work
Main Menu > File > Save
Export Mesh

35
Fig.5.12 Export mesh file
Main Menu > File > Export > Mesh...
Save the file as wedge.msh.
Make sure that the Export 2d Mesh option is selected.
5.5 FLUENT:
This software solve/iterate the problem by importing the file which was exported by
GAMBIT to define the problem in FLUENT. It contains various options to cover all sorts of
problems starting from simple fluid flow problem to complex fluid structural interaction
(FSI).
Fluent will require following data to setup the solution for problem.
 Solution method/model
 Material properties like density, viscosity, thermal conductivity, specific heat or
internal energy etc.
 Boundary and operating conditions for given problem so that given physics can be
revealed.
 Initial conditions and no. of iterations required to converge the solution. This serves
as initial step of algorithm
 After applying/specifying above data, FLUENT is ready for solution process. It will
start iterating till the solution gets converged

36
 In FLUENT, solution is performed by use of various solvers which serves as base of
any solution. FLUENT contains various solvers SIMPLE, SIMPLEC, k-epsilon
models or turbulence model etc.
 After solution is performed; post processing is done for reviewing the results of
solutions to analyze the given problem.
Problem Identification and Pre-Processing
1. Define your modeling goals.
2. Identify the domain you will model.
3. Design and create the grid.
Solver Execution
4. Set up the numerical model.
5. Compute and monitor the solution.
Post-Processing
6. Examine the results.
7. Consider revisions to the model.
Fig.5.13 Gambit- Fluent

37
5.6 Initial Settings
(Double Click)Setup in the Workbench Project Page.
When the FLUENT Launcher appears change options to "Double Precision", and then
click OK as shown below.The Double Precision option is used to select the double-precision
solver. In the double-precision solver, each floating point number is represented using 64 bits
in contrast to the single-precision solver which uses 32 bits. The extra bits increase not only
the precision, but also the range of magnitudes that can be represented. The downside of
using double precision is that it requires more memory.
Fig.5.14 Fluent launcher
5.6.1 Problem Setup - General
Now, FLUENT should open. We will begin setting up some options for the solver. In the left
hand window (in what I will call the Outline window), under Problem Setup, select General.
The only option we need to change here is the type of solver. In the Solver window,
select Density-Based.
Fig 5.15 General

38
5.6.2 Models
In the outline window, click Models. We will need to utilize the energy equation in order to
solve this simulation. Under Models highlight Energy - Off and click Edit.... Now,
the Energy window will launch. Check the box next to Energy Equation and hit OK. Doing
this turns on the energy equation.
We also need to change the type of viscosity model. Select Viscous - Laminar and press OK
Fig.5.16 Models
5.6.3 Materials
In the Outline window, highlight Materials. In the Materials window, highlight Fluid, and
click Create/Edit.... this will launch theCreate/Edit Materials window; here we can specify
the properties of the fluid. Set the Density to Ideal Gas, the Specific Heat to1006.43,
the Molecular Weight to 28.966. When you have updated these fields, press Change/Create.
Fig.5.17 Materials

39
5.7 Boundary Conditions
In the Outline window, select Boundary Conditions. We will now specify each boundary
condition for the simulation.
5.7.1 Pressure far-field
In the Boundary Conditions window, select Pressure far-field. Use the drop-down menu to
change the Type to pressure-far-field. You will be asked to confirm the change, and do so by
pressing OK. Next, a dialogue box will open with some parameters we need to specify.
Change the Gauge Pressure (Pascal) to 101325, and Mach Number to 2.95.
Fig.5.18 Pressure far-field
Also, select the Thermal tab, and ensure that the temperature correctly defaulted to 300 K.
When you are finished, press OK.
5.8 Operating Conditions
In the Boundary Conditions window, select the Operating Conditions button. Change
the Gauge Pressure to 0. Then press OK
Fig.5.19 operating conditions

40
It is important to check the operating conditions. When setting the density in materials to
ideal gas, FLUENT calculates the density using the absolute pressure. However, the pressure
we specify is the gauge pressure, not the absolute pressure. FLUENT will use the absolute
pressure to compute the density therefore if we do not set the operating pressure to 0 our
density will be incorrect for the flow field.
5.8.1 Reference Values
In the Outline window, select Reference Values. Change the Compute From parameter
to Pressure far-field. Check that the values are accurate. The reference values are used when
calculating the non-dimensional results such as the drag coefficient.
Fig.5.20 Reference values
5.9 Solution Methods
In the Outline window, select Solution Methods to open the Solution Methods window.
Under Spatial Discretization, ensure that the option under Flow First Order Upwind is
selected.

41
Fig.5.21 Solution methods
5.10 Solution Controls
In the Outline window, select Solution Controls to open the Solution Controls window.
Ensure that the Courant Number is set to 0.01.
The Courant number can be considered a non dimensionalized timestep. The density-based
solver obtains the steady-state solution by starting with the initial guess and marching in
pseudo-time until convergence is obtained. The Courant number controls the time step the
solver uses. The larger it is, the faster the solution will converge but it will not be very stable
and can diverge. The smaller it is, the slower it is to reach convergence but the solution is
much more stable.
Fig.5.22 Solution controls

42
5.11 Monitors
In the Outline window, click Monitors to open the Monitors window. In
the Monitors window, select Residuals - Print,Plot and press Edit.... This will open
the Residual Monitors window. We want to change the convergence criteria for our solution.
Under Equationand to the right of Continuity, change the Absolute Criteria to 0.001. Repeat
for x-velocity, y-velocity, and energy, then press OK.
Fig.5.23 Residual monitors
5.12 Solution Initialization
In the Outline window, select Solution Initialization. We need to make an "Initial Guess" to
the solution so FLUENT can iterate to find the final solution. In the Solution
Initialization window, select Standard Initialization, then under Compute from,
select Pressure far-field from the drop down box. Check to see that the values that generate
match our inputted values, and then press Initialize.
Fig.5.24 Solution initialization

43
Run Calculation
In the Outline window, select Run Calculation. Change the Number of Iterations to 20000.
Double click Calculate to run the calculation. It should a few minutes to solve. After the
calculation is complete, save the project. Do not close FLUENT.
Fig.5.25 Run calculation
RESULTS AND COMPARISON:
Fig.5.26 Scaled Residuals with Slot

44
Fig.5.27 Scaled Residuals without Slot
Fig.5.28 (a) Total Temperature Contour with slot

45
Fig.5.28 (b) Total Temperature Contour without slot
Fig.5.29 (a) Static Temperature Contour with slot

46
Fig.5.29 (b) Static Temperature Contour without slot
Fig.5.30 (a) Static Pressure Contour with slot

47
Fig.5.30 (b) Static Pressure Contour without slot
Fig.5.31 (a) Velocity Contour with slot

48
Fig.5.31 (b) Velocity Contour without slot
Fig.5.32 Enthalpy Contour with slot

49
Fig.5.33 (a) Density Contour with slot
Fig.5.33 (b) Density Contour without slot

50
Fig.5.34 (a) Plot of Total Pressure with Slot
Fig.5.34 (b) Plot of Total Pressure without Slot

51
Fig.5.35 (a) Plot of Static Pressure with Slot
Fig.5.35 (b) Plot of Static Pressure without Slot

52
Fig.5.36 (a) Plot of Static Temperature with Slot
Fig.5.36 (b) Plot of Static Temperature without Slot

53
Fig.5.37 (a) Plot of Total Temperature with Slot
Fig.5.37 (b) Plot of Total Temperature without Slot

54
Fig.5.38 (a) Plot of velocity with Slot
Fig.5.38 (b) Plot of velocity without Slot

55
Fig.5.39 (a) Plot of Enthalpy with Slot
Fig.5.39 (b) Plot of Enthalpy without Slot

56
Fig.5.40 (a) Plot of Density with Slot
Fig.5.40 (b) Plot of Density without Slot

57
CHAPTER6
Conclusion
Application of a theory which is a solution of the momentum and continuity
equations for inlets with zero angle cowls estimated quite closely the
experimental pressure recoveries obtained. The use of throat bleed in the
vicinity of the center body shoulder controlled flow separation sufficiently to
allow attachment of the internal lip shock. By use of this bleed slot the
performance of the supersonic inlet is varied. The comparison of supersonic
inlet without slot with the supersonic inlet with slot data indicates that with
proper boundary-layer control no severe shock - boundary-layer interaction
losses were incurred by the use of the extreme cowl-lip angles at this Mach
number.

58
Chapter 7
References
1. J.Thomas Anderson
Details of the Geometry and Operation of the SR-71 Mixed Compression Inlet
By Technical Fellow Emeritus Lockheed Martin Skunk Works.
2. Images of F35.
3. D.Kliche ch.Mundt E H Hrischel.
The hypersonic Mach number independence principle in the case of viscous flow
4. MATTINGLY
Element of Propulsion- Gas turbine and rockets
5. YUE LianJie, XU XianKun & CHANG XinYu.
Theoretical analysis of effects of boundary layer bleed on scramjet thrust SDF
Aerospace and Aerodynamics Corner
6. Hill and Peterson
Mechanics and thermodynamics of propulsion
7. Simon trapier Philippe duveau Sebastian deck
Experimental study of supersonic inlet buzz.

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